Rotary compressors or turbines



A ril 10, 1962 e. M. LEWIS 3,029,011

ROTARY COMPRESSORS 0R TURBINES Filed Oct. 8, 1956 2 Sheets-Sheet 1 April10, 1962 G. M. LEWIS 3,029,011

ROTARY COMPRESSORS OR TURBINES Filed Oct. 8, 1956 2 Sheets-Sheet 2United States Patent 3,029,011 ROTARY CMPRESUES R TlJRlElNES GordonMantis Lewis, Bristol, England, assignor, by mesne assignments, toBristol Siddclcy Engines Limited,

Bristol, England, a British company Filed Oct. 8, 1956, Ser. No. 614,546Claims priority, application Great Britain Get. 13, 1955 1 Claim. (Q1.230-114) This invention relates to axial flow rotary compressors andturbines using a gaseous medium (hereinafter referred to as gas), and inthe case of such compressors has as its main object to provide a meansof stabilising the operation of the compressor when the speed ofrotation ditlers substantially from the optimum operating speed, usuallyreferred to as the designed speed. The invention is applicable to singleand multi-stage compressors as well as single and multi-stage turbinesand more especially, although not exclusively, to the compressors andturbines of axial flow gas turbine engines. As is well known, in amulti-stage compressor or turbine each stage comprises a ring of movingblades and a ring of stationary blades.

With multi-stage, axial flow, rotary compressors at speeds of rotationbelow the designed speed, when the pressure rise per stage is less thanthe designed pressure rise, the reduction in axial velocity of the maingas flow at the entry to the compressor results in the moving blades ofthe initial stages of the compressor operating at an increasing angle ofincidence until, when the speed of rotation has been reducedsufficiently, the blades become stalled. A blade ring does not howeverstall simultaneously over its whole area, but initially only certainparts stall, the axial velocity falling substantially in these parts sothat the main flow is concentrated in the unstalled parts and asufficient axial velocity is maintained in the unstalled parts to delayfurther their stalling. The rate at which the sta ling spreads, and theextent of the blading affected is a measure of the severity of thestalling.

One object of the present invention is to provide means whereby theaxial velocity of the gas flow past a blade ring approaching a stalledor partially stalled condition may be increased whereby the stalling isdelayed and the speed range increased. This eifect is achieved,according to the invention, by reducing the effective cross-sectionalarea of the gas flow passage upstream of the blade ring, and theinvention can be used for this purpose in 'urbines as well ascompressors.

According to the present invention in an axial flow, rotary gascompressor or turbine having. a plurality of blade rings, there isprovided an annular slot or an annular series of slots in a wall of thegas flow passage of the compressor or turbine on the upstream side ofand adjacent a blade ring, and supply means for supplying the slot orslots with gas under pressure always greater than the pressure in thegas flow passage in the region of the slot or slots, the slot or slotsbeing arranged to inject gas so supplied to them into the gas flowpassage with a component of motion which is directed upstream againstthe main gas flow through the gas flow passage, to reduce the effectiveflow area of the gas flow passage in the region of the slot or slots andthereby to increase the velocity of the gas flow through the gas flowpassage downstream of the slot or slots.

In the case of a compressor, the stalling or partial stalllag of theearly stage blading may, or may not precipitate full compressor surge,depending on the severity of the stall in the early stages. In caseswhere the early stage stall is sufficiently violent, the compressor isunable to operate with these stages in their stalled state over acertain range of r.p.m., and a kink in the surge line results, the surgeline being that line on a plot of compression 'ice ratio against massgas flow at various constant rotational speeds which divides the area ofthe plot corresponding to possible operating conditions of thecompressor from the area in which stable operation cannot take place.v

The existence of such a kink in the surge line has the disadvantage thatthe compressor has to be operated in general at a condition further fromthe surge line, and therefore at less efliciency, in order to avoidcrossing the surge line at the kink.

It has been observed that in most axial flow compressors stalling of thefirst ring of moving blades commences in zones spaced around thecircumference of the ring at the tips of the blades, such zones movingaround to produce what is known as a rotating stall pattern, and thatcompressors behaving in this way often exhibit the above describedundesirable kink in their surge lines. On the other hand, the firstmoving blade rings of compressors having certain features of design, forexample rather large blade tip clearances, tend to stall at the tipsevenly all round the circumference of the blade ring, and it has beenobserved that with these compressors the kink in the surge line is muchless marked.

Unfortunately the features of design which tend to produce thischaracteristic behaviour also usually tend to reduce the efliciency ofthe compressor at its designed speed.

According to a feature of the present invention, however, it has beenfound that by positioning said slot or slots immediately upstream of theblade tips of a rotor blade ring, the interference with the main gasflow caused by injection of the gas through said slot or slots promotesstalling at the blade tips of said rotor blade ring. By adopting thisfeature of the invention therefore circumferentially distributed tipstalling may be promoted in any moving blade ring of the compressor, andit has been found possible by this means to reduce the undesirable kinkin the surge line of the compressor.

Three embodiments of the present invention will now be described, merelyby Way of example and with reference to the accompanying drawings,whereof FIGURE 1 is a partial cross sectional elevation of onemulti-stage axial flow rotary gas compressor according to the invention,

FIGURE 2 is a partial cross sectional elevation of another constructionof'multi-stage axial flow rotary gas compressor according to theinvention, and

FIGURE 3 is a partial cross sectional elevation of yet anotherconstruction of multi-stage axial flow rotary gas compressor accordingto the invention.

Referring to FIGURE 1, the moving or rotor blades of the compressor, thefirst ring of which is indicated at 1.0, are carried on a drum 11mounted for rotation inside a stationary casing 12 which casing carriesrings of stationary blades, the first ring 13 of which is in an annularair intake 14 of the compressor upstream of the first rotor.

blade ring, and forms an entry guide blade ring to the compressor, andthe subsequent rings 15 of which are positioned alternately with therotor blade rings in well known manner.

In the construction being described the longest blades of the compressorare at the entry end and are represented by the entry guide blades 13,while the shortest blades are at the discharge end of the compressor,the blades being made progressively shorter from the entry to thedischarge end.

According to the invention gas under pressure is tapped off from anintermediate stage of the compressor or from the compressor delivery andpassed through conduit means, part of which is indicated at 19, to anannular gallery 16 in the stationary casing 12, the annular gallerysurrounding the ring of entry guide blades 13. The

gallery 16 communicates by a number of passages 17 with an annular slot18 formed in the casing 12 immediately upstream of the blade tips of theblades of the first ring of rotor blades. Instead of the annular slot 18there may be provided an annular series of slots in the casing 12immediately upstream of the ring 10.

The annular slot 18 may be of a nozzle formation and is arranged toinject the gas which is supplied to it with a component of motion whichis directed upstream against the main gas flow through the gas fiowpassage of the compressor. In the example at present being described theslot 13 is arranged to inject gas at an angle of about 60 to a planenormal to the rotational axis of the rotor 11.

The flow of tapped-off gas supplied to the slot 18 through the conduitmeans is controlled by a valve (not shown) in the conduit meanspreviously described and the flow of tapped oil? gas may amount to about2% of the total gas fiow through the compressor.

The flow control valve is preferably of the kind which gas turbineengine with increase in the delivery of fuel to the engine.

In the example at present being described, assuming that there is noinjection of gas from the slot 18, when the speed of the compressor isreduced sufiiciently below the designed speed, there is a tendency forthe gas flow in the main gas flow passage of the compressor to bereversed over the blade tips of the first rotor blade ring and to formvortices. As already stated however, this usually occurs in patches anda rotating pattern is formed.

By injecting gas through the slot 18 energy is added to the vorticesadjacent the blade tips of the rotor blade ring 16 so that thesevortices join up and become stabilised in the form of a single toroidalvortex. The injection eirect ensures that all the blade tips in thefirst rotor blade ring are substantially encircled by the toroidalvortex and results in the development or" a sta led zone of relativelyuniform annular shape and this in turn results in a smooth transition tothe fully stalled condition. In addition, the injected air creates adead zone, occupied by the toroidal vortex, with the result that theeffective cross-sectional area of the main gas flow passage of thecompressor at the first rotor blade ring is reduced. By reducing theeffective cross-sectional area of the main gas flow passage in thisregion the velocity of gas flow over the blades of the first rotor bladering and the blade rings downstream of the first rotor blade ring isincreased, and this has the effect of improving the effective matchingof the inlet stages of the compressor with the rest of the compressorstages with the result that a Wider flow range is possible beforesurging occurs.

In the example now being described therefore the present invention maybe seen to operate in two Ways (a) by effectively improving the matchingof the compressor stages at speeds below the designed speed and reducingthe compressor capacity, thereby moving the surge line in the directionof lower flow, and

(b) by promoting and stabilising the partly stalled condition in theinitial stage of the compressor, resulting in removal of the kink in thesurge line.

With multi-stage axial flow rotary gas compressors and turbines gas maybe injected into the main gas flow passage at more than one blade ring,but in the case of a compressor the best results appear to occur whenthe injection is made so as to promote the formation of blade tipvortices at the rotor blade ring having the longest blades as justdescribed. It will be appreciated however that the operation of theinvention as described at (a) above may be achieved by arranging anannular slot such as 13 on the upstream side of and adjacent the entryguide blades 13 so that injection of gas through the slot causes theelfective cross-sectional area of the gas flow passage upstream of theentry guide blade ring to be reduced, and thereby the velocity of gasflow over the blading of the compressor to be increased. The slot isposh tioned within about a blade length of the guide blades, the bladelength being that of the guide blades. An example of an arrangement inwhich a slot such as 18 is positioned on the upstream side and adjacentthe entry guide blade ring of a compressor is later described withreference to FIGURES 2 and 3.

la the case where the injection of air is to take place immediatelyupstream of a rotor blade ring it is possible to apply the inventionalso or alternatively to the shortest rotor blades of the compressordescribed with reference to F GURE 1 so that the tips of these bladesare stalled at speeds of rotation of the compressor above the designedspced when the pressure rise per stage is greater than the designpressure rise. Such a case might arise where maximum efficiency of thecompressor is designed to occur at partial loading.

Referring now to FIGURE 2 in which the same reference numerals are usedto indicate parts already described with reference to FIGURE 1, there isprovided in this case, as previously mentioned, an annular slot 26formed in structure 21 defining the inner Wall of the annular air intake14 to the compressor. The structure 21 is formed with an annular gallery22 which is communicated by conduit means part of which is shown at 23with an interrnediatc stage or the delivery end of the compressor, valvemeans being provided in the conduit means which valve may be of t- .ekind previously described.

The slot which is of nozzle formation, is located on the upstream sideof and adjacent the ring of entry guide blades 13. and is arranged toinject gas supplied to the gallery 22 into the air intake 14 of thecompressor with a component of motion upstream against the main gas flowthrough the air intake, in the present example at an angle of about 60to the rotational axis of the rotor 11.

As in the previous case the annular slot 20 may if desired be replacedby an annular series of slots, and the slot 2% is provided in additionto the slot 13 which is positioned as previously described.

During operation of the compressor at the speed below its designed speedthe injection of gas through the slot 20 reduces the effectivecross-sectional area of the annular air intake to the compressor in theregion of the slot with the result that the velocity of gas flow overthe binding of the compressor is increased, the matching of thecompressor stages thereby being improved and the compressor capacitybeing reduced thereby moving the surge line in the direction of lowerflow.

In the case of an axial-flow rotary gas compressor where a slot such as2% as described with reference to FIGURE 2 only is provided, the slotmay be formed in the outer casing structure on the upstream side of andadjacent the entry guide blades to the compressor. Such a constructionis shown in FIGURE 3 to which reference will now be made. In this casean annular slot 24 having a nozzle formation is formed in the outercasing structure 25 of the compressor which outer casing structuresupports a ring of entry guide blades 26 at the entry end of thecompressor and also subsequent stator rings of the compressor stages inwell known manner.

The slot 24 is supplied with gas from an annular gallery 39 formed inthe outer casing structure, the gallery 3% being communicated with anintermediate stage or the delivery end of the compressor by conduitmeans (not shown) the conduit means including a valve to control theflow of tapped oil gas through the conduit means which valve may be ofthe kind hereinbefore described.

The slot 24 is, as before, arranged to inject gas with a component ofmotion which is directed upstream against the main gas flow through theair intake 27 to the compressor, in the present example at an angle of60 to the axis of the compressor rotor.

In order further to increase the restriction of the effectivecross-sectional area of the air intake of the com pressor described withreference to FIGURE 3 by the injection of gas into the intake a furtherannular slot may be formed in the structure 31 defining the inner wallof the intake 27, the slot in the structure 31 lying in the same radialplane as the slot 24, and being arranged to inject gas supplied to itinto the intake with a component of motion directed upstream as before,means being provided to supply the second slot with gas under pressurefrom an intermediate stage or the delivery end of the compressor.

The range over which the valve controlling the supply of gas to anyparticular slot or annular series of slots is arranged to be opendepends on the position of the compressor operating line without gasinjection through the slot in relation to the kink in the surge line ofthe compressor. The valve is arranged to open over that range of theoperating line of the compressor without gas injection which approachesthe kink in the surge line. This range may be an initial range or anintermediate range.

As previously indicated the present invention may be applied to an axialflow rotary gas turbine. Where the invention is applied to such aturbine it is preferred that the injection of gas into the main gas flowpassage of the turbine be arranged to occur between a stator blade ringand the next downstream rotor blade ring. The gas to be injected ispreferably tapped from the main gas flow passage of the turbine or a gasturbine engine of which the turbine forms a part and is, of course,injected at a pressure higher than that of the gas flow at the place ofinjection with a component of motion which is directed upstream againstthe main gas fiow through the gas flow passage of the turbine. When thetapping is not feasible or desirable, gas may be taken from anothersource but must be injected under a suflicient pressure to deflect themain gas How to the desired extent.

By injecting gas into the gas flow passage of an axial flow rotary gasturbine in the manner described, the effective cross-sectional area ofthe main gas flow passage of the turbine is reduced in the region atwhich the injection takes place, and the velocity of the gas flow overthe turbine blading downstream of the region is increased. This has theeffect of reducing the capacity of the turbine and may be used, forexample, to improve the matching of a rotary axial flow gas turbine of agas turbine engine with a compressor of the engine when the turbine isrunning at a speed below its designed speed.

In FIGURES 1 and 2 of the drawings, the resultant flow pattern of thegases is indicated at 14a and 18a. At 18a is indicated the fluid barrierset up by the gas injected into the gas flow passage 14 through the slotformation 18 and at 14a is indicated the resultant flow of gas throughthe passage 14 which is produced by the barrier 18a.

In FIGURE 3 of the drawings, the resultant flow pattern of the gases isindicated at 24a and 27a. At 24a is indicated the fluid barrier set upby the gas injected into the gas flow passage 27 through the slotformation 24 and at 27a is indicated the resultant fiow of gas throughthe passage 27 which is produced by the barrier 24a In the drawings theblades have been shown as nonshrouded blades. The invention is howeverapplicable to axial flow, rotary compressors and turbines havingshrouded as well as non-shrouded blades.

I claim:

A rotary axial flow bladed power conversion machine comprising inner andouter walls defining an annular gas flow passage, and a plurality ofblade rings contained in said passage, there being in one of said walls,immediately upstream of the leading edge of the blades of a rotor bladering, an annular slot formation which is coaxial with said passage andhas a discharge opening slanted in an upstream direction and transverseto the annular gas flow passage and which serves for injecting into saidgas flow passage during operation of the machine gas having an upstreamcomponent of velocity greater than the downstream component of velocityof the gas in said gas flow passage and under sufficient pressure tosubstantially deflect the main gas flow, the slot formation beingdefined at least in part by bounding surfaces which have a directiveeffect on the injected gas, said surfaces directing the injected gaswith a component of motion upstream in the gas flow passage whereby theeffective flow area of the gas flow passage in the region of the slotformation is reduced and the velocity of the gas flow through the gasflow passage downstream of the slot formation and through the blade ringis increased.

References Cited in the file of this patent UNITED STATES PATENTS701,500 Olsson June 3, 1902 1,111,498 Rotter Sept. 22, 1914 2,404,275Clark et al. July 16, 1946 2,418,801 Baumann Apr. 8, 1947 2,599,470Meyer June 3, 1952 2,660,366 Klein et a1 Nov. 24, 1953 2,685,429 AuyerAug. '3, 1954 2,718,349 Wilde Sept. 20, 1955 2,749,027 Stalker June 5,1956 2,763,427 Lindsey Sept. 18, 1956 2,763,984 Kadosch et al. Sept. 25,1956 2,864,236 Toure et al Dec. 16, 1958' 2,957,306 Attinello Oct. 25,1960 2,958,456 Forshaw Nov. 1, 1960 FOREIGN PATENTS 504,214 GreatBritain Apr. 21, 1939 507,316 Italy Dec. 29, 1954 611,447 Great BritainOct. 29, 1948 745,630 Great Britain Feb. 29, 1956.

757,496 Great Britain Sept. 19, 1956 963,540 France Jan. 4, 1950-

